Propellant Mixture Research Articles

A study of an air-breathing electrodeless plasma thruster discharge

Plasma chemistry of main Earth atmospheric components in VLEOs is implemented in a hybrid 2D axisymmetric simulation code to assess the air-breathing concept in an electrodeless plasma thruster. Relevant electron-heavy species collisions for diatomic molecules, and atom associative wall recombination into molecules are included. Simulations are run by injecting 1 mg/s of Xe, N2 and O independently for powers between 10 and 3000 W. The performances and trends of plasma response for N2 and O are similar to Xe but displaced to higher powers. Since they have lighter elementary masses, a higher plasma density is generated and more electrons need to be heated. At optimum power, the thrust efficiency for N2 and O surpasses that of Xe, which is caused by the excess of neutral re-ionization and the associated inelastic and wall losses. Additional simulations are run injecting 50/50 of N2/O to study the thruster operation for propellant mixtures, and the performances are found to be linear combinations of those of each propellant in the absence of collisions between heavy species. Injection of O2 is also studied for the impact of the possible associative recombination of O at the intake walls, and the performances are found similar to those of O due to the strong molecular dissociation inside the thruster.

Characterization of New Minimum Signature Solid Propellant Based on Ammonium Perchlorate Synergetic with Sorbitol and Magnesium Metal

Since the introduction of the concept of insensitive weapons, rocket motor designers must now ensure that their vulnerability is as low as possible. One of the inherent characteristics of a sugar-based propellant is the obvious plume coming out from the nozzle as a result of chemical reactions. The present work characterises a novel minimum signature composite solid propellant based on Ammonium perchlorate/Sorbitol with magnesium metal as a fuel and chlorine scavenger. The elemental stoichiometric coefficient and oxygen balance for the propellant mixture were calculated to find the optimum ratio between the oxidizer and fuel. The solid propellant of AP was prepared by mixing it with melted sorbitol and casting it into the cylindrical hollow tube using the freestanding method. The burning rate for five samples was measured at atmospheric pressure. The morphology was characterized by a scanning electron microscope. Calculation results reveal that ammonium perchlorate at 0.77% weight produced the best elemental stoichiometric coefficient and oxygen balance. The sample with the additive metal showed a significant increase in the burning rate. In all formulation samples of magnesium and ferrous oxide, the maximum burning rate at 0.126 cm/s showed that the metal additive and the propellant worked synergistically with each other. Results from SEM indicate that ammonium perchlorate spherical particles were uniformly dispersed within sorbitol. The addition of ferrousoxide increases the interaction between the component, and at the same time it gives the oxidizer the ability to interact more efficiently with the fuel and improve the physical strength of the solid propellant. This research is a minor step for Malaysia to become a top competitor in the aerospace industry in the future.

Air-breathing rotating detonation engine supplied with liquid kerosene: propulsive performance and combustion stability

Experimental results are presented for a rotating detonation engine supplied with liquid kerosene and preheated air without liquid or gaseous additions to the propellant mixture. Various combustion modes for the generic combustor geometry design were observed—from deflagration, through pulsed combustion and high-frequency instabilities, to stable detonation propagation. Attention was paid to detonation stability (if present), its characteristics, and the propulsive performance of the combustor with a focus on specific thrust and pressure gain through thrust and outlet total pressure measurement. These parameters measured for the observed modes were compared. The stability of the detonation combustion proved not to be critical to achieve high performance of the combustion chamber. For example, high performance was achieved for combustion modes with high-frequency instabilities.

Impacts of metal oxide crystalline structure on the decomposition of solid propellants under combustion heating rates

Ammonium perchlorate (AP) has been the oxidizer of choice for composite solid propellants for decades and has been the object of decomposition studies for safety monitoring. Typically, studies perform differential scanning calorimetry (DSC) or thermogravimetric analysis (TGA) to show how metal oxides (MO) commonly incorporated into propellants alter AP decomposition rates and completeness of reaction. Most past decomposition work studies temperatures below the crystal phase transition (240°C) from orthorhombic phase to cubic, which impose internal stresses within the lattice of AP particles. Phase change induces a partial decomposition, which does not follow a shrinking core behavior; instead, it develops a network of pores up to a few microns in size throughout. During low-temperature decomposition, particles lose approximately 30-40 % of their mass. Simultaneous DSC/TGA (STA) use low heating rates for combustion environments but offer information on MO's catalytic and electron transport modifications. This work demonstrates how the crystalline morphology MO additives mechanistically alter the combustion properties of composite propellants by identifying modes for promoting decomposition in AP and HTPB. MO morphology determines the electron structure of the molecule, which sets the band gap properties of the particles. This study evaluated different morphologies and a range of MOs to identify the most active MO for decomposing A.P. Kissinger analysis was applied using STA data at heating rates of 10-, 20-, and 30-°C/min revealing significant shift in the high-temperature decomposition with a range of metal oxides. Higher heating rates were evaluated using CO2 laser ignition to identify the time to first gas using the MO that offers the best heat absorption, such as the aluminum oxides. This higher heating rate was found do more accurately represent the changes in the combustion rates in the final propellant mixture. Additionally, it was shown that electron transport additives like CuO show the most significant impact on combustion, and thermal absorbers offer the lowest impact on AP/HTPB combustion. This understanding offers a new approach to propellant design not previously presented and suggests future testing for tailoring the use of metal oxides in propellants.

Optimal Liquid Engine Architecture by Performance-Cooling Tradeoff Analysis

We perform physics-based theoretical analyses on the tradeoff between the overall performance of specific impulse and the thermal protection ability of a coolant liquid film to identify the optimal configuration of a liquid rocket engine. A bipropellant thruster is set as the target, which uses a hypergolic propellant mixture of nitrogen tetroxide as the oxidant and monomethyl hydrazine as the fuel. By considering the practical nonuniform distribution of the local mixture ratio produced in the thrust chamber, the maximum specific impulse is achieved when the fuel and oxidizer spray widths become identical, allowing for the specification of the diameter of the impinging-type injector. The heat balance between the convective heat transfer from the combustion gas and the latent heat of the three-dimensionally wavy liquid film provides the optimal diameter of the combustion chamber. The longest liquid film is found to be achieved when half of the initial film is entrained by the fast gas, correspondingly mitigating the heat transfer area due to the liquid film waviness. We successfully demonstrate the optimal architecture of the liquid engine based on the tradeoff, in which an improvement of specific impulse by 1 s shortens the film length by 2 mm.

Experimental and numerical investigation on agglomeration of aluminum during combustion process of aluminized composite propellant

The addition of aluminum particles can effectively improve the energy characteristics of solid composite propellants in propulsion systems. However, these aluminum particles affecting propellant mixture viscosity, ignition delay time, burning time and specific impulse could also lead to undesirable effects. Therefore, it is a general practice that in the early stage of propellant formulation design that the particle size of condensed phase products is optimized/minimized through formulation adjustment to reduce its negative impact on the engine. In practice, it means that a costly manpower and materials are needed to obtain the size distribution of condensed phase products during the combustion process of propellants via experimental tests. Thus, there is a strong industrial need for a reasonable and feasible simulation method of agglomeration process during the combustion process of propellants. In this work, we propose and examine the implementation of molecular dynamics and a random packing algorithm on creating numerically models of aluminized composite propellants. Considering the pyrolysis of ammonium perchlorate (AP) and the combustion process of aluminum, we develop a more accurate agglomeration model basing on the classical Jackson model. Our proposed methods and simulations models are validated by our experimental results. It confirms that the present work provides a prediction tool for simulating and improving the agglomeration process, thus aiming at optimizing solid propellant formula.

Rock splitting with pyrotechnic compositions and secondary propellants

In some cases of resources extraction, as well as in construction at industrial and urbanized territories, commercial explosives are not safe enough for the surroundings with regard to the generated fly-rocks, air-blast, toxic fumes, seismic waves and vibrations. The main reasons for these harmful impacts of explosion are the velocity and the mechanism of the chemical reaction of explosive decomposition. The authors shifted the focus of their researches from detonating explosives to high-speed combusting energetic materials. The production of low explosive non-detonating mixtures from long-term stored smokeless gunpowders and ammonium nitrate prills in different configurations, and popular pyrolant compositions were studied. The samples of different cartridge casings, filled with non-detonating propellant mixtures or pyrotechnic compositions were examined by two methods - for velocity of propagation and by field tests. It was made investigations for application of waste single-base-powders and double-base-powders for obtaining non-detonating explosive cartridges, suitable in dimension stone mining, as well as in blasting activities at tender and complicated conditions.

IGNITION OF NITROMETHANE-BASED PROPELLANT MIXTURES

Hydrazine is a commonly used propellant in spacecraft and satellite propulsion, but it is toxic as well as carcinogenic, and consequently the high costs associated with its handling have led the propulsion community to seek alternatives. Our research group is exploring the use of nitromethane, a low-toxicity substance and a staple laboratory solvent, as the primary ingredient for a green monopropellant. However, nitromethane's ignition is challenging, requiring a certain level of pressure for stable combustion. To improve ignition and combustion behavior, we screened additives including transition metal acetylacetonates and other organometallic compounds. Gas-tight vessels in a differential scanning calorimetry (DSC) device and single-droplet combustion tests were used to measure the influence of additives on the exotherm onset and peak in nitromethane thermal decomposition. The screening showed that some transition-metal acetylacetonates and other organometallic compounds effectively enhance thermal decomposition in sealed-vessel DSC measurements. Ferrocene had the most significant influence: addition of 2 wt. % of this substance to nitromethane caused a decomposition peak shift from 384°C to 280°C. In single-droplet combustion tests in an oxygen-deprived atmosphere at 3 bar, the addition of ferrocene caused a 40% reduction of the droplet ignition delay time, combined with a 20% higher combustion rate compared to neat nitromethane. This makes ferrocene a strong candidate for the use as an ignition and combustion catalyst in nitromethane-based propellants.

Liquid resistivity of pharmaceutical propellants using novel resistivity cell

Metered-dose inhalers employ propellants to produce pharmaceutical aerosols for treating respiratory conditions like asthma. In the liquid phase, the DC volume resistivity of pharmaceutical propellants, including R134a, R152a, and R227ea, was studied at saturation pressures and room temperature (not vapour phase). These measurements are essential for industries like refrigerants. Aerosols from metered dose inhalers (MDIs) with these propellants become electrically charged, affecting medicament deposition in lung. The resistivity was measured using a novel concentric cylinder-type capacitance cell designed in-house. The resistivity for the propellants (R134a, R152a, and R227ea) was found to be 3.02 × 1010 Ωm, 2.37 × 109 Ωm and 1.31 × 1010 Ωm, respectively. The electrical resistivity data obtained was found to be at least two orders of magnitude higher than the limited data available in the literature. Challenges in the resistivity cell’s development and performance are discussed, with a focus on various propellants and their mixtures with ethanol and moisture concentrations. The resistivity of propellant mixtures containing moisture concentrations ranging from 5 to 500 ppm and ethanol concentrations ranging between 1000 and 125,000 ppm was determined. The resistivity was tested across 10-min and 1-h periods and was performed in accordance with the contemporary IEC 60247 standard.

Point-to-Point Efficient Grafting Improved Compatibility of Boron-Based Fuel-Rich Propellants.

Due to the high mass calorific value, boron powder is widely used in energetic material systems such as solid propellants and ignition powders. Especially, boron powder has very broad application prospects in the field of fuel-rich propellants. However, due to the cross-linking reaction between the boric acid contained on the surface of boron powder and an adhesive called hydroxylated polybutadiene (HTPB), the viscosity of the propellant mixture system increases sharply, which seriously affects the preparation of the propellant. In this paper, "click chemistry" is intended to be used to graft the functional groups on the surface of boron powder to reduce the viscosity of the boron powder and HTPB in the initial mixing stage. In addition, a rheometer was used to test the viscosity of the boron powder and the HTPB system. The test results showed that compared to the viscosity of the raw boron powder system at 24.1 Pa·s, the mixing termination viscosity of the grafted sample was 17.1 Pa·s, a decrease of 29.0%.

Перспективи використання азотовмісних однокомпонентних ракетних па-лив

The goal of this work is to analyze the possibility of using existing monopropellant compositions based on aqueous solutions of high-energy nitrogen-containing substances as the main propellant for low-thrust engines, for example, for meteorological rockets, for upper-stage engines, and in spacecraft control engine systems. This paper presents an approach that considers the selection and justification of ingredients based on renewable energy sources, the analysis being carried out primarily from standpoint of the availability of propellant components and their safety and energy efficiency. It is proposed that the energy of unitary reducing agent – oxidizer chemical propellants (energy-saturated compositions) be used as an alternative source. The development of nonhydrocarbon nitrogen-containing alternative energy sources with the possibility of their conversion and accumulation into the planetary nitrogen, oxygen, and water cycles is an urgent problem. The paper presents detailed information on propellant mixtures of nitrogen-containing substances as oxidizers and considers a number of reducing agents, such as alcohols, amides, etc. in composition with high-energy additives (aluminum, magnesium). The calculated results obtained meet the objectives and demonstrate that the compositions considered can be used as the main propellant for low-thrust engines. The advantages of the new propellant technology: availability, a low cost, produceability, environmental friendliness, a relatively low toxicity, and, primarily, a simpler design of the propulsion system and launch equipment. The proposed propellant composition, which is under test, is planned for use in the sustainer engines of ultralight suborbital rockets with the possibility of further development to an orbital rocket system.

Crystal Structures of the Commonly Used Plasticizers Ethylene Glycol Dinitrate, Diethylene Glycol Dinitrate, and Triethylene Glycol Dinitrate─Characterization and Discussion of Their Structural Behavior

Ethylene glycol dinitrate (EGDN), diethylene glycol dinitrate (DEGDN), and triethylene glycol dinitrate (TEGDN) are used as plasticizers in propellant mixtures and have been known for more than 100 years. Despite the industrial application and the long history of these compounds, the crystal structures of all three compounds, which are liquids at room temperature, have not been determined. Therefore, in this work, the crystal structures were examined by low-temperature X-ray diffraction, and thus, the bonding properties and crystal packing in the solid state could be compared and discussed for the first time. Furthermore, the compounds were characterized by nuclear magnetic resonance spectroscopy and infrared spectroscopy. The thermal properties were investigated by DTA (differential thermal analysis) measurements, and the experimental vapor pressures were measured with the chromatography-assisted transpiration method. The densities at room temperature (for the liquids) were measured by gas pycnometry, and the energetic properties were calculated using the EXPLO5 code. Moreover, a more in-depth analysis of the different sensitivities through discussion of the Hirshfeld surfaces, created based on their crystal structures, was performed to compare their sensitivities in the solid state.

Optical visualization of hypergolic burning spray structure using blue light spectrum

In this study, we focused on a novel methodology involving the optical visualization of a hypergolic burning spray. A hypergolic propellant, comprising 95 wt% H2O2 and an amine-based fuel, was injected using pentad impingement for implementing the hypergolic reacting flow. The intense luminosity of hypergolic flame was successfully suppressed using blue light spectrum illumination and bandpass optical filtering. The hypergolic burning spray structures can be classified into reactive stream separation, normal combustion, and oxidizer partial consumption based on the dynamic interaction of impinging jets. The majority of dispersed spray propellant was consumed in developing a hypergolic diffusion flame in the vicinity of a fuel-to-oxidizer dynamic pressure ratio (DR) of 0.83, a design condition of optimum mixing. Either off-design combustion of excessive oxidizer (DR ≪ 0.83) or fuel dynamic pressure (DR ≫ 0.83) caused deformation of the burning spray structure. The results indicated that the burning spray structure is significantly affected by the dynamic pressure ratio of impinging jets. The burning rate of the propellant mixture is expected to vary significantly with respect to the hypergolic burning structure. A simple optical visualization of the reacting flow via the blue light spectrum enhanced the qualitative understanding of hypergolic combustion.

Synthesis of two novel neutral polymeric bonding agents to enhance the mechanical properties of composite solid propellants.

Currently, only a few bonding agents can be utilized efficiently in nitramine filler material-based solid rocket energetic binder systems. Herein, we demonstrate the synthesis and specific characterization of two kinds of neutral polymeric bonding agents (NPBA-1 and NPBA-2) and their application in composite propellants consisting of the nitramine octogen (HMX) and glycidyl azide polymer (GAP). The as-obtained NPBAs were well-coated on the surface of HMX and RDX due to their functionalized groups, and they significantly affected the viscosity of the uncured propellant mixtures and possessed obviously enhanced mechanical properties in the cured AP/HMX/GAP propellant mixtures, even at low concentrations (down to 0.001 wt% of the whole propellant). In addition, because of the existence of an epoxy group and no hydroxyl functionalities, NPBA-2 exhibited improved mechanical strength and glass transition temperature as compared to NPBA-1, which has plenty of reactive hydroxyl groups. The as-synthesized epoxy-modified NPBAs are essential for obtaining NEPE propellants with high bonding and mechanical properties.

COATING VITON ON FLAKE ALUMINUM AND ITS EFFECTS ON PERFORMANCE OF THE SOLID ROCKET MOTOR

Aluminum powder is used as an additive in composite propellants to increase the specific impulse. However, the use of aluminum increases the two-phase losses and slag formation in the rocket motor. This study uses Viton as a coating for the aluminum powder for enhancing its reactivity to overcome these drawbacks. Viton-coated aluminum powder exhibited an enhanced reactivity over the uncoated powder. This Viton-coated aluminum demonstrated an increase of around 25% in the burn rates compared to uncoated aluminum when used in the composite solid propellant. The end mix viscosity of the propellant mixture remains constant with the prolonged mixing with the use of Viton compared to an increased viscosity observed with polytetrafluoroethylene as an additive.

Parameters Influencing the Characteristic Exhaust Velocity of a Nitrous Oxide/Ethene Green Propellant

Currently, several green propellants are under investigation to replace the toxic and carcinogenic monopropellant hydrazine (). Beside alternatives as ammonium dinitramide-based propellants, hydroxylammonium nitrate-based mixtures, or high concentrated hydrogen peroxide, mixtures of fuels with nitrous oxide () called HyNOx or NOFBX represent promising green propellants. Compared to classical , mixtures offer a higher performance ( and ), no or low toxicity, and low propellant costs. To investigate the handling and performance of a propellant mixture consisting of and ethene, hot gas combustion tests with an experimental combustor were conducted. This paper summarizes the results of 134 combustion tests conducted with the premixed propellant injected in gaseous state. Calculated and measured performances ( and ) depending on mixture ratio, characteristic combustion chamber length , and chamber pressure are shown and discussed. Furthermore, the residence timescales of the propellant mixture inside the combustion chamber are normalized by the chemical timescale of the combustion process. Thus, combustion efficiencies depending on dimensionless Damköhler numbers are derived.

Timescale-Based Frozen Nonadiabatic Flamelet Combustion Modeling for Rocket Engine Thrust Chambers

The present research work introduces a novel combustion model formulation based on the flamelet concept suitable to investigate characteristic flow conditions in rocket engine thrust chambers. The new approach extends the validity domain of existing flamelet models to fully cover the observed flow conditions by incorporating kinetic rate effects into a state-of-the-art nonadiabatic flamelet manifold. This enables a transition between the fluid dynamic and reaction-rate-dominated combustion regimes, depending on the local Damköhler number of the numerical solution field. The approach is investigated based on a representative rocket engine test case focusing on highly nonadiabatic and de Laval nozzle flow conditions for the propellant combination methane-oxygen. Covering a range of propellant mixture ratios associated with significant variations of the underlying chemical effects, three flamelet combustion-modeling approaches of different manifold complexity are compared to high-fidelity laminar finite rate results based on the full chemical reaction mechanism. The obtained data show that the proposed model formulation is able to deliver a notably higher accuracy in the numerical prediction compared to the existing modeling approaches while retaining high computational efficiency to meet industrial design requirements.

Effects of 2CL-20/HMX cocrystals on the thermal decomposition behavior and combustion properties of polyether solid propellants

Herein, the effects of cocrystals of two nitroamine explosives, namely hexanitrohexaazaisowurtzitane (CL-20) and 1,3,5,7-tetranitro-1,3,5,7-tetrazoctane (HMX), which are denoted as CHC, on the combustion properties, heat release, and safety profiles of polyether solid propellants were systematically analyzed. The peak decomposition temperature of CHC was found to be 244.8 °C, which is lower than those of its component CL-20 and HMX. However, the thermal stability of CHC was similar to that of CL-20, and the thermal decomposition behavior of CHC was significantly different from that of a physical mixture of CL-20 and HMX. In addition, similar to CL-20, CHC was not completely decomposed upon heating, despite the decomposition activation energy of CHC (322.7 kJ·mol−1) being similar to that of HMX (322.4 kJ·mol−1). In contrast, the thermal decomposition behaviors of CHC-based propellants differed from those of solid propellants that were based on physical mixtures of CL-20 and HMX. During thermal decomposition, CHC did not decompose into CL-20 and HMX, and the heat released from the CHC-based propellants was more concentrated than that released from the other propellant mixture examined. Furthermore, the rate of combustion of CHC-based propellants fell between those of the CL-20- and HMX-based propellants at all the tested pressures. The pressure exponents of combustion rates of the examined CHC-based propellants were slightly lower than those of the CL-20- and HMX-based propellants, and the combustion rates and their corresponding pressure exponents decreased with a decrease in the CHC particle size. Moreover, the explosive heat of CHC-based propellant was between those of the CL-20-based and HMX-based propellants, and similar to that of a CL-20/HMX mixture-based propellant. Finally, the safety performance of CHC-based propellant was almost identical to that of the CL-20/HMX mixture.

Power efficiency estimation of an inductive plasma generator using propellant mixtures of oxygen, carbon-dioxide and argon

A promising aspect of inductive plasma generators for space propulsion applications is their electrodeless nature that leads to a high propellant compatibility and opens the door to in-situ resource utilisation propellants. To further develop inductive plasma generators as a space propulsion technology it is important to understand the relationship between system inputs (e.g. input power and propellant material) and system outputs (e.g. power efficiencies) to aid design and control. In this work, methodologies are developed for non-intrusively assessing the inductive coupling from antenna current measurements in terms of the ‘instantaneous ignition power’, the ‘inductive duty cycle’ and the ‘inductive frequency shift’. Experimental results are presented for IPG7, a high-power (up to 180 kW) 5.5-turn helical antenna inductive plasma generator. Potential in-situ resource utilisation materials oxygen and carbon-dioxide, and their respective mixtures with argon, are assessed in terms of their inductive coupling characteristics and resulting power efficiencies. Oxygen is shown to be an attractive propellant material, especially when supplemented with argon. In particular, high thermal efficiency (∼84%) can be achieved at high input powers. A basis is provided for understanding the power efficiencies of an inductive plasma generator in terms of non-intrusive antenna current measurements, to aid power supply sizing and to develop monitoring and control techniques for thruster applications.

Effects of Ammonium Perchlorate Particle Size, Ratio, and Total Contents on the Properties of a Composite Solid Propellant

AbstractIn this study, ammonium perchlorates (APs) with two or more different particle sizes were mixed to optimize the viscosity of the propellant slurry and the burn rate of the resulting solid composite propellants at various mixing ratios. First, to investigate the characteristics of bimodal AP, APs with 200 and 6 μm particles were used. The experimental results showed that the viscosity of the propellant mixture was excellent when the AP‐200 content, i. e., the AP with 200 μm particles, ranged from 55 to 70 wt%, and the burn rate increased with the increasing AP‐6 content. To investigate the characteristics of trimodal AP, three types of APs with 400, 200, and 6 μm particles were used. The findings revealed that the mixing ratio of the three types of APs most favorable to the rheological properties was approximately 1 : 1 : 1. In addition, although adjusting the ratio between large and small particles was desirable for changing the burn rate, varying the ratio between AP‐400 and AP‐200 did not significantly change the burn rate. Adjusting the AP‐6 ratio effectively changed the burn rate. The method of adjusting the AP contents, i. e., the solid contents of propellant, can effectively alter the viscosity and burn rate of propellant. When the total AP content was reduced, the burn rate decreased, but the overall propellant performance degraded. However, the performance parameters of the propellant, such as its density and specific impulse, could be improved by adding solid fuels, such as aluminum.